Method and apparatus for actively controlling fuel flow to a mixer assembly of a gas turbine engine combustor

ABSTRACT

An apparatus for actively controlling fuel flow from a fuel pump to a mixer assembly of a gas turbine engine combustor, where the mixer assembly includes a pilot mixer and a main mixer. The pilot mixer further includes an annular pilot housing having a hollow interior, a primary fuel injector mounted in the pilot housing and adapted for dispensing droplets of fuel to the hollow interior of the pilot housing, a plurality of axial swirlers positioned upstream from the primary fuel injector. The fuel flow control apparatus further includes: at least one sensor for detecting dynamic pressure in the combustor; a fuel nozzle; and, a system for controlling fuel flow supplied by the fuel nozzle through the valves. The fuel nozzle includes: a feed strip with a plurality of circuits for providing fuel to the pilot mixer and the main mixer; and, a plurality of valves associated with the fuel nozzle and in flow communication with the feed strip thereof. The control system activates the valves in accordance with signals received from the pressure sensor.

BACKGROUND OF THE INVENTION

The present invention relates to a staged combustion system in which theproduction of undesirable combustion product components is minimizedover the engine operating regime and, more particularly, to a method andapparatus for actively controlling fuel flow to a mixer assembly havinga pilot mixer with a primary fuel injector and secondary fuel injectionports.

Modern day emphasis on minimizing the production and discharge of gasesthat contribute to smog and to other undesirable environmentalconditions, particularly those gases that are emitted from internalcombustion engines, have led to different gas turbine engine combustordesigns that have been developed in an effort to reduce the productionand discharge of such undesirable combustion product components. Otherfactors that influence combustor design are the desires of users of gasturbine engines for efficient, low cost operation, which translates intoa need for reduced fuel consumption while at the same time maintainingor even increasing engine output. As a consequence, important designcriteria for aircraft gas turbine engine combustion systems includeprovision for high combustion temperatures, in order to provide highthermal efficiency under a variety of engine operating conditions, aswell as the minimization of undesirable combustion conditions thatcontribute to the emission of particulates, and to the emission ofundesirable gases, and to the emission of combustion products that areprecursors to the formation of photochemical smog.

Various governmental regulatory bodies have established emission limitsfor acceptable levels of unburned hydrocarbons (HC), carbon monoxide(CO), and oxides of nitrogen (NOx), which have been identified as theprimary contributors to the generation of undesirable atmosphericconditions. Therefore, different combustor designs have been developedto meet those criteria. For example, one way in which the problem ofminimizing the emission of undesirable gas turbine engine combustionproducts has been attacked is the provision of staged combustion. Inthat arrangement, a combustor is provided in which a first stage burneris utilized for low speed and low power conditions to more closelycontrol the character of the combustion products. A combination of firststage and second stage burners is provided for higher power outletconditions while attempting to maintain the combustion products withinthe emissions limits. It will be appreciated that balancing theoperation of the first and second stage burners to allow efficientthermal operation of the engine, while simultaneously minimizing theproduction of undesirable combustion products, is difficult to achieve.In that regard, operating at low combustion temperatures to lower theemissions of NOx, can also result in incomplete or partially incompletecombustion, which can lead to the production of excessive amounts of HCand CO, in addition to producing lower power output and lower thermalefficiency. High combustion temperature, on the other hand, althoughimproving thermal efficiency and lowering the amount of HC and CO, oftenresults in a higher output of NOx.

Another way that has been proposed to minimize the production of thoseundesirable combustion product components is to provide for moreeffective intermixing of the injected fuel and the combustion air. Inthat regard, numerous mixer designs have been proposed over the years toimprove the mixing of the fuel and air. In this way, burning occursuniformly over the entire mixture and reduces the level of HC and COthat result from incomplete combustion. Even with improved mixing,however, higher levels of undesirable NOx are formed under high powerconditions when the flame temperatures are high.

One mixer design that has been utilized is known as a twin annularpremixing swirler (TAPS), which is disclosed in the following U.S. Pat.Nos. 6,354,072; 6,363,726; 6,367,262; 6,381,964; 6,389,815; 6,418,726;6,453,660; 6,484,489; and, 6,865,889. It will be understood that theTAPS mixer assembly includes a pilot mixer which is supplied with fuelduring the entire engine operating cycle and a main mixer which issupplied with fuel only during increased power conditions of the engineoperating cycle. While improvements in the main mixer of the assemblyduring high power conditions (i.e., take-off and climb) are disclosed inpatent applications having Ser. Nos. 11/188,596, 11/188,598, and11/188,470, modification of the pilot mixer is desired to improveoperability across other portions of the engine's operating envelope(i.e., idle, approach and cruise) while maintaining combustionefficiency.

In order to provide increased functionality and flexibility, the pilotmixer in a TAPS type mixer assembly has been developed and is disclosedin a patent application entitled “Pilot Mixer For Mixer Assembly Of AGas Turbine Engine Combustor Having A Primary Fuel Injector And APlurality Of Secondary Fuel Injection Ports.” This patent application,having Ser. No. 11/365,428, is owned by the assignee of the presentapplication and hereby incorporated by reference. While the '428application is concerned with the physical embodiments of the pilotmixer, it will be appreciated that an apparatus and method is desiredwhich is able to actively control fuel flow to such pilot mixer, as wellas the overall mixer assembly containing it.

It is well known that lean, premix combustion requires operation closeto the lean-blow out boundary in order to minimize emissions. Therefore,it is desired that the onset of a lean blow out event be recognized sothat operation of the combustor can be adjusted and lean blow outavoided. In addition, the mixing of air and fuel must be extremelyeffective to achieve low emissions. To enhance such mixing, pulsing thefuel to the injectors at a high frequency would also be desirable.

It has also been found that lean, premix combustion often results inhigh dynamic pressure levels in the combustor. The combustion dynamicsis a result of interaction between heat release from combusting thefuel-air mixture and pressure oscillations in the chamber. Such dynamicpressures may result in high cycle fatigue and can damage combustorparts. While the effects of dynamic pressures on the combustor have beencountered previously, this has generally involved the provision of highbandwidth fuel or air actuation to reduce the pressure levels associatedwith acoustic modes of the combustor.

Thus, there is a need to provide a gas turbine engine combustor in whichthe production of undesirable combustion product components is minimizedover a wide range of engine operating conditions. Accordingly, it isdesired that the pilot mixer of a nested combustor arrangement bemodified to include a primary fuel injector and a plurality of secondaryfuel injection ports. It is also desired that an active control systemand process be provided which enhances operation of such mixer assemblyby identifying and countering the onset of a lean blow out condition, aswell as an unacceptable level of dynamic pressure experienced in thecombustor.

BRIEF SUMMARY OF THE INVENTION

In a first exemplary embodiment of the invention, an apparatus foractively controlling fuel flow from a fuel pump to a mixer assembly of agas turbine engine combustor is disclosed, where the mixer assemblyincludes a pilot mixer and a main mixer. The pilot mixer furtherincludes an annular pilot housing having a hollow interior, a primaryfuel injector mounted in the pilot housing and adapted for dispensingdroplets of fuel to the hollow interior of the pilot housing, and aplurality of axial swirlers positioned upstream from the primary fuelinjector. The fuel flow control apparatus further includes: at least onesensor for detecting dynamic pressure in the combustor; a fuel nozzle;and, a system for actively controlling fuel flow supplied to the pilotmixer and the main mixer of the mixer assembly by the fuel nozzle. Thefuel nozzle further includes: a feed strip with a plurality of circuitsfor providing fuel to the pilot mixer and the main mixer; and, aplurality of valves associated with the fuel nozzle and in flowcommunication with the feed strip thereof. The control system activatesthe valves in accordance with signals received from the pressure sensor.

In a second exemplary embodiment of the invention, an apparatus foractively controlling fuel flow from a fuel pump to a mixer assembly of agas turbine engine combustor is disclosed, where the mixer assemblyincludes a pilot mixer and a main mixer. The pilot mixer furtherincludes an annular pilot housing having a hollow interior, a primaryfuel injector mounted in the pilot housing and adapted for dispensingdroplets of fuel to the hollow interior of the pilot housing, aplurality of axial swirlers positioned upstream from the primary fuelinjector, and a plurality of secondary fuel injection ports forintroducing fuel into the hollow interior of the pilot housing. The fuelflow control apparatus further includes: at least one sensor fordetecting dynamic pressure in the combustor; a fuel nozzle; and, asystem for actively controlling fuel flow supplied to the pilot mixerand the main mixer of the mixer assembly by the fuel nozzle. The fuelnozzle further includes: a feed strip with a plurality of circuits forproviding fuel to the primary fuel injector of the pilot mixer, thesecondary fuel injection ports of the pilot mixer, and the main mixer;and, a plurality of valves associated with the fuel nozzle and in flowcommunication with the feed strip thereof. The control system activatesthe valves in accordance with signals received from the pressure sensor.

In a third exemplary embodiment of the invention, a method of activelycontrolling fuel flow from a fuel pump to a mixer assembly of a gasturbine engine combustor is disclosed, the mixer assembly including apilot mixer and a main mixer, wherein the pilot mixer further includesan annular pilot housing having a hollow interior and a primary fuelinjector mounted in the pilot housing and adapted for dispensingdroplets of fuel to the hollow interior of the pilot housing. The methodincludes the following steps: continuously sensing dynamic pressure in acombustion chamber of the combustor; determining whether an amplitude ofthe sensed dynamic pressure in the combustion chamber is greater than apredetermined amount; and, signaling a fuel nozzle to provide fuel in aspecified manner to the pilot mixer when the pressure amplitude isgreater than the predetermined amount.

In a fourth exemplary embodiment of the invention, a method of activelycontrolling fuel flow from a fuel pump to a mixer assembly of a gasturbine engine combustor is disclosed, the mixer assembly including apilot mixer and a main mixer, wherein the pilot mixer further includesan annular pilot housing having a hollow interior, a primary fuelinjector mounted in the pilot housing and adapted for dispensingdroplets of fuel to the hollow interior of the pilot housing, and aplurality of secondary fuel injection ports for introducing fuel intothe hollow interior of the pilot housing. The method includes thefollowing steps: continuously sensing dynamic pressure in a combustionchamber of the combustor; determining whether an amplitude of the senseddynamic pressure in the combustion chamber is greater than apredetermined amount; and, signaling a fuel nozzle to provide fuel in aspecified manner to the secondary fuel injection ports of the pilotmixer when the pressure amplitude is greater than the predeterminedamount.

In a fifth exemplary embodiment of the invention, a method of activelycontrolling fuel flow from a fuel pump to a mixer assembly of a gasturbine engine combustor is disclosed, the mixer assembly including apilot mixer and a main mixer, wherein the pilot mixer further includesan annular pilot housing having a hollow interior and a primary fuelinjector mounted in the pilot housing and adapted for dispensingdroplets of fuel to the hollow interior of the pilot housing. The methodincludes the following steps: continuously sensing dynamic pressure in acombustion chamber of the combustor; determining whether a frequency ofthe sensed dynamic pressure in the combustion chamber is within apredetermined range; and, signaling a fuel nozzle to provide fuel in aspecified manner to the pilot mixer when the pressure frequency iswithin the predetermined range.

In a sixth exemplary embodiment of the invention, a method of activelycontrolling fuel flow from a fuel pump to a mixer assembly of a gasturbine engine combustor is disclosed, the mixer assembly including apilot mixer and a main mixer, wherein the pilot mixer further includesan annular pilot housing having a hollow interior, a primary fuelinjector mounted in the pilot housing and adapted for dispensingdroplets of fuel to the hollow interior of the pilot housing, and aplurality of secondary fuel injection ports for introducing fuel intothe hollow interior of the pilot housing. The method includes thefollowing steps: continuously sensing dynamic pressure in a combustionchamber of the combustor; determining whether a frequency of the senseddynamic pressure in the combustion chamber is within a predeterminedrange; and, signaling a fuel nozzle to provide fuel in a specifiedmanner to the secondary fuel injection ports of the pilot mixer when thepressure frequency is within the predetermined range.

In a seventh exemplary embodiment of the invention, a method of activelycontrolling fuel flow from a fuel pump to a mixer assembly of a gasturbine engine combustor during a plurality of operational stages isdisclosed, the mixer assembly including a pilot mixer and a main mixer,wherein the pilot mixer further includes an annular pilot housing havinga hollow interior, a primary fuel injector mounted in the pilot housingand adapted for dispensing droplets of fuel to the hollow interior ofthe pilot housing, and a plurality of secondary fuel injection ports forintroducing fuel into the hollow interior of the pilot housing. Themethod includes the following steps: supplying fuel only to the primaryfuel injector and the secondary fuel injection ports of the pilot mixerduring a first fueling mode; supplying fuel to the pilot mixer and themain mixer in a first specified amount during a second fueling mode;and, supplying fuel to the pilot mixer and the main mixer in a secondspecified amount during a third fueling mode.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a diagrammatic view of a high bypass turbofan gas turbineengine;

FIG. 2 is a longitudinal, cross-sectional view of a gas turbine enginecombustor having a staged arrangement;

FIG. 3 is an enlarged, cross-sectional view of the mixer assemblydepicted in FIG. 2;

FIG. 4 is a cross-sectional view of a fuel nozzle assembly and the mixerassembly depicted in FIGS. 2 and 3;

FIG. 5 is a block diagram of a system for providing fuel flow to themixer assembly depicted in FIGS. 2 and 3;

FIG. 6 is a schematic diagram of a system for actively controlling fuelflow through the fuel nozzle assembly depicted in FIG. 4;

FIG. 7 is a flow diagram depicting operational steps involved in aprocess for actively controlling fuel flow through the fuel nozzleassembly depicted in FIG. 4 to the mixer assembly depicted in FIGS. 2and 3; and,

FIG. 8 is a diagrammatic view of how fuel flow is provided to the mixerassembly depicted in FIGS. 2 and 3 during specified stages of engineoperation.

DETAILED DESCRIPTION OF THE INVENTION

Referring now to the drawings in detail, wherein identical numeralsindicate the same elements throughout the figures, FIG. 1 depicts indiagrammatic form an exemplary gas turbine engine 10 (high bypass type)utilized with aircraft having a longitudinal or axial centerline axis 12therethrough for reference purposes. Engine 10 preferably includes acore gas turbine engine generally identified by numeral 14 and a fansection 16 positioned upstream thereof. Core engine 14 typicallyincludes a generally tubular outer casing 18 that defines an annularinlet 20. Outer casing 18 further encloses and supports a boostercompressor 22 for raising the pressure of the air that enters coreengine 14 to a first pressure level. A high pressure, multi-stage,axial-flow compressor 24 receives pressurized air from booster 22 andfurther increases the pressure of the air. The pressurized air flows toa combustor 26, where fuel is injected into the pressurized air streamto raise the temperature and energy level of the pressurized air. Thehigh energy combustion products flow from combustor 26 to a first (highpressure) turbine 28 for driving high pressure compressor 24 through afirst (high pressure) drive shaft 30, and then to a second (lowpressure) turbine 32 for driving booster compressor 22 and fan section16 through a second (low pressure) drive shaft 34 that is coaxial withfirst drive shaft 30. After driving each of turbines 28 and 32, thecombustion products leave core engine 14 through an exhaust nozzle 36 toprovide propulsive jet thrust.

Fan section 16 includes a rotatable, axial-flow fan rotor 38 that issurrounded by an annular fan casing 40. It will be appreciated that fancasing 40 is supported from core engine 14 by a plurality ofsubstantially radially-extending, circumferentially-spaced outlet guidevanes 42. In this way, fan casing 40 encloses fan rotor 38 and fan rotorblades 44. Downstream section 46 of fan casing 40 extends over an outerportion of core engine 14 to define a secondary, or bypass, airflowconduit 48 that provides additional propulsive jet thrust.

From a flow standpoint, it will be appreciated that an initial air flow,represented by arrow 50, enters gas turbine engine 10 through an inlet52 to fan casing 40. Air flow 50 passes through fan blades 44 and splitsinto a first compressed air flow (represented by arrow 54) that movesthrough conduit 48 and a second compressed air flow (represented byarrow 56) which enters booster compressor 22. The pressure of secondcompressed air flow 56 is increased and enters high pressure compressor24, as represented by arrow 58. After mixing with fuel and beingcombusted in combustor 26, combustion products 60 exit combustor 26 andflow through first turbine 28. Combustion products 60 then flow throughsecond turbine 32 and exit exhaust nozzle 36 to provide thrust for gasturbine engine 10.

As best seen in FIG. 2, combustor 26 includes an annular combustionchamber 62 that is coaxial with longitudinal axis 12, as well as aninlet 64 and an outlet 66. As noted above, combustor 26 receives anannular stream of pressurized air from a high pressure compressordischarge outlet 69. A portion of this compressor discharge air flowsinto a mixing assembly 67, where fuel is also injected from a fuelnozzle 68 to mix with the air and form a fuel-air mixture that isprovided to combustion chamber 62 for combustion. Ignition of thefuel-air mixture is accomplished by a suitable igniter (not shown), andthe resulting combustion gases 60 flow in an axial direction toward andinto an annular, first stage turbine nozzle 72. Nozzle 72 is defined byan annular flow channel that includes a plurality of radially-extending,circularly-spaced nozzle vanes 74 that turn the gases so that they flowangularly and impinge upon the first stage turbine blades of firstturbine 28. As shown in FIG. 1, first turbine 28 preferably rotates highpressure compressor 24 via first drive shaft 30. Low pressure turbine 32preferably drives booster compressor 24 and fan rotor 38 via seconddrive shaft 34.

Combustion chamber 62 is housed within engine outer casing 18 and isdefined by an annular combustor outer liner 76 and a radially-inwardlypositioned annular combustor inner liner 78. The arrows in FIG. 2 showthe directions in which compressor discharge air flows within combustor26. As shown, part of the air flows over the outermost surface of outerliner 76, part flows into combustion chamber 62, and part flows over theinnermost surface of inner liner 78.

Contrary to previous designs, it is preferred that outer and innerliners 76 and 78, respectively, not be provided with a plurality ofdilution openings to allow additional air to enter combustion chamber 62for completion of the combustion process before the combustion productsenter turbine nozzle 72. This is in accordance with a patent applicationentitled “High Pressure Gas Turbine Engine Having Reduced Emissions” andhaving Ser. No. 11/188,483, which is also owned by the assignee of thepresent invention. It will be understood, however, that outer liner 76and inner liner 78 preferably include a plurality of smaller,circularly-spaced cooling air apertures (not shown) for allowing some ofthe air that flows along the outermost surfaces thereof to flow into theinterior of combustion chamber 62. Those inwardly-directed air flowspass along the inner surfaces of outer and inner liners 76 and 78 thatface the interior of combustion chamber 62 so that a film of cooling airis provided therealong.

It will be understood that a plurality of axially-extending mixingassemblies 67 are disposed in a circular array at the upstream end ofcombustor 26 and extend into inlet 64 of annular combustion chamber 62.It will be seen that an annular dome plate 80 extends inwardly andforwardly to define an upstream end of combustion chamber 62 and has aplurality of circumferentially spaced openings formed therein forreceiving mixing assemblies 67. For their part, upstream portions ofeach of inner and outer liners 76 and 78, respectively, are spaced fromeach other in a radial direction and define an outer cowl 82 and aninner cowl 84. The spacing between the forwardmost ends of outer andinner cowls 82 and 84 defines combustion chamber inlet 64 to provide anopening to allow compressor discharge air to enter combustion chamber62.

A mixing assembly 100 in accordance with one embodiment of the presentinvention is shown in FIG. 3. Mixing assembly 100 preferably includes apilot mixer 102, a main mixer 104, and a cavity 106 positionedtherebetween. More specifically, it will be seen that pilot mixer 102preferably includes an annular pilot housing 108 having a hollowinterior, as well as a primary fuel injector 110 mounted in housing 108and adapted for dispensing droplets of fuel to the hollow interior ofpilot housing 108. Further, pilot mixer 102 preferably includes a firstswirler 112 located at a radially inner position adjacent primary fuelinjector 110, a second swirler 114 located at a radially outer positionfrom first swirler 112, and a splitter 116 positioned therebetween. Asshown, splitter 116 extends downstream of primary fuel injector 110 toform a venturi 118 at a downstream portion. It will be understood thatfirst and second pilot swirlers 112 and 114 are generally orientedparallel to a centerline axis 120 through mixing assembly 100 andinclude a plurality of vanes for swirling air traveling therethrough.Fuel and air are provided to pilot mixer 102 at all times during theengine operating cycle so that a primary combustion zone 122 is producedwithin a central portion of combustion chamber 62 (see FIG. 2).

Main mixer 104 further includes an annular main housing 124 radiallysurrounding pilot housing 108 and defining an annular cavity 126, aplurality of fuel injection ports 128 which introduce fuel into annularcavity 126, and a swirler arrangement identified generally by numeral130. Swirler arrangement 130 may be configured in any of several ways,as seen in a patent application entitled “Mixer Assembly For CombustorOf A Gas Turbine Engine Having A Plurality Of Counter-Rotating Swirlers”having Ser. No. 11/188,596 and a patent application entitled “SwirlerArrangement For Mixer Assembly Of A Gas Turbine Engine Combustor HavingShaped Passages” having Ser. No. 11/188,595, both of which are assignedto the owner of the present invention. It will be seen in FIG. 3,however, that swirler arrangement 130 preferably includes at least afirst swirler 144 positioned upstream from fuel injection ports 128. Asshown, first swirler 144 is preferably oriented substantially radiallyto centerline axis 120 through mixer assembly 100. It will be noted thatfirst swirler 144 includes a plurality of vanes 150 for swirling the airflowing therebetween. Since vanes 150 are substantially uniformly spacedcircumferentially, a plurality of substantially uniform passages aredefined between adjacent vanes 150. It will further be understood thatswirler 144 may include vanes having different configurations so as toshape the passages in a desirable manner, as disclosed in the '595patent application identified hereinabove.

Swirler arrangement 130 also is shown as including a second swirler 146positioned upstream from fuel injection ports 128 and preferablyoriented substantially parallel to centerline axis 120. Second swirler146 further includes a plurality of vanes 152 for swirling the airflowing therebetween. Although vanes 152 are shown as beingsubstantially uniformly spaced circumferentially, thereby defining aplurality of substantially uniform passages therebetween, such vanes 152may also have different configurations so as to shape the passages in adesirable manner.

Cavity 106, as stated above, is located between pilot mixer 102 and mainmixer 104 and contains a first fuel manifold 107 in flow communicationwith a fuel supply. In particular, a centerbody outer shell 140 forms anouter surface and an aft surface of cavity 106, with pilot housing 108providing an inner surface thereof. Fuel injection ports 128 are in flowcommunication with fuel manifold 107 and spaced circumferentially aroundcenterbody outer shell 140. As seen in FIG. 3, fuel injection ports 128are preferably positioned so that fuel is provided in an upstream end ofannular cavity 126.

When fuel is provided to main mixer 104, an annular, secondarycombustion zone 198 is provided in combustion chamber 62 that isradially outwardly spaced from and concentrically surrounds primarycombustion zone 122. Depending upon the size of gas turbine engine 10,as many as twenty or so mixer assemblies 100 can be disposed in acircular array at inlet 64 of combustion chamber 62.

As seen in FIG. 3, pilot mixer 102 also preferably includes a pluralityof spaced secondary fuel injection ports 134, whereby fuel is alsointroduced into hollow interior of pilot housing 108. It will beappreciated that secondary fuel injection ports 134 are preferablyspaced circumferentially about pilot housing 108 within a designatedplane 136 intersecting centerline axis 120 through mixing assembly 100.

While plane 136, in which secondary fuel injection ports 134 lie, isshown as being located in a flared portion 138 of pilot housing 108downstream of splitter 116, it will be understood that a planecontaining such secondary fuel injection ports 134 may be located atapproximately a downstream end of splitter 116 or even upstream thereof.Indeed, the axial length of splitter 116 may be altered so that itsrelationship with the location of secondary fuel injection ports 134could change.

Similarly, plane 136 is depicted as being oriented substantiallyperpendicular to centerline axis 120, but secondary fuel injection ports134 may be positioned so that plane 136 is skewed so as to be angledeither upstream or downstream as desired. Further, regardless of theaxial position or orientation of plane 136 containing secondary fuelinjection ports 134, each such secondary fuel injection port 134 mayindividually be oriented substantially perpendicular to centerline axis120, oriented upstream at an acute angle, or oriented downstream at anobtuse angle.

It will further be seen that secondary fuel injection ports 134 of pilotmixer 102 preferably are in flow communication with a second fuelmanifold 109, which also is preferably located within cavity 106. Fuelis typically injected into the hollow portion of pilot housing 108 bysecondary fuel injection ports 134 upon the occurrence of a specifiedevent (e.g., a designated cycle point for gas turbine engine 10, whencompressor discharge air 58 is a designated temperature, etc.).Depending upon the requirements of a specific condition, fuel isinjected through secondary fuel injection ports 134 at a rate greaterthan, less than or substantially the same as fuel injected throughprimary fuel injector 110. Of course, this presumes that fuel will beprovided by primary fuel injector 110 at all times, but there may beoccasions when it is preferable to provide fuel to pilot mixer 102 onlythrough secondary fuel injection ports 134.

In this way, pilot mixer 102 has greater flexibility during operationacross the lower power conditions (i.e., idle, approach and cruise). Inparticular, it will be appreciated that pilot mixer 102 is able to powergas turbine engine 10 up to approximately 30% of maximum thrust whenfuel is provided solely to primary fuel injector 110. By comparison,pilot mixer 102 is able to power gas turbine engine 10 up toapproximately 70% of maximum thrust when fuel is provided to secondaryfuel injection ports 134 as well.

In order to promote the desired fuel spray into the hollow interior ofpilot housing 108, it is preferred that a passage 142 surround eachsecondary fuel injection port 134 of pilot mixer 102. Each passage 142is in flow communication with compressed air via a supply 154 providedin cavity 106. This air is provided to facilitate injection of the fuelspray into pilot housing 108 instead of being forced along an innersurface 156 thereof. This may further be enhanced by providing a swirler158 within each passage 142 which provides a swirl to the air injectedaround the fuel spray.

It is also preferred that vanes of outer pilot swirler 114 be configuredso that air passing therethrough is directed at least somewhat towardinner surface 156 of pilot housing 108. In this way, such air is betterable to interact with fuel provided by secondary fuel injection ports134. Accordingly, such vanes are preferably angled at approximately 30°to about 60° with respect to centerline axis 120. In this way, a flareangle 160 of pilot housing 108 is approximated.

Considering the addition of secondary fuel injection ports 134 in pilotmixer 102, it will be appreciated that the flow rate of air therethroughis preferably maintained at a rate of approximately 10% to approximately30%. Further, such secondary injection ports 134 assist in reducing theemissions produced by mixer assembly 100 during the operation of gasturbine engine 10. In particular, combustor 26 is able to operate onlywith fuel being supplied to pilot mixer 102 for a greater time period.Also, it has been found that providing more fuel at a radially outerlocation of pilot mixer 102 is desirable.

It will further be seen in FIGS. 4-7 that an apparatus and method forcontrolling fuel flow to mixer assembly 100 is provided. With respect tofuel nozzle 68, it will be appreciated that it is configured similar tothat shown and described in U.S. Pat. No. 6,955,040 to Myers, Jr. etal., which is hereby incorporated herein. More specifically, it will beseen that fuel nozzle 68 includes a housing 174 located at an outerradial location which contains a plurality of valves, a nozzle support176 which extends between valve housing 174 and mixer assembly 100, anda macrolaminate feed strip 178 positioned within nozzle support 176.Feed strip 178 further includes a first circuit 180 for supplying fuelto a fuel tube 132 (which is in flow communication with primary fuelinjector 110 of pilot mixer 102), a second circuit 182 for supplyingfuel to fuel manifold 109 (which is in flow communication with secondaryfuel injection ports 134 of pilot mixer 102), and a third circuit 183for supplying fuel to fuel manifold 107 (which is in flow communicationwith fuel injection ports 128 of main mixer 104).

In order to better understand the manner in which fuel is supplied tomixer assembly 100, a block diagram of an overall fuel flow controlsystem 200 is depicted in FIG. 5. As seen therein, system 200 includes afuel pump 202, whereby a fuel supply (not shown) in flow communicationtherewith provides fuel to each fuel nozzle 68 positioned around annularcombustor 26. A fuel nozzle control 204 is provided for each fuel nozzle68 in order to generally control the valves within housing 174 andtherefore the amount of fuel provided by circuits 180, 182 and 183. Fuelnozzle control 204 interfaces with fuel pump 202 and receives signals208 from a full authority digital engine control (FADEC) 206 tocoordinate the proper fueling mode of pilot and main mixers 102 and 104depending upon the current stage of operational cycle for gas turbineengine 10. This will be explained in more detail herein with respect toFIG. 8.

It will be appreciated that staging valves 184, 186 and 188, which areassociated with circuits 180, 182 and 183, respectively, are activatedaccording to a signal 210 provided by fuel nozzle control 204. Fuel isthen permitted to flow through first circuit 180, second circuit 182,and third circuit 183 within feed strip 178 of each fuel nozzle 68according to the positioning of staging valves 184, 186 and 188. In thisway, fuel is either provided in the desired amount to primary fuelinjector 110 of pilot mixer 102, secondary fuel injection ports 134 ofpilot mixer 102, and fuel injection ports 128 of main mixer 104 of eachmixer assembly 100.

In order to pulse the fuel in first, second, and/or third circuits 180,182 and 183, a second separate control signal 212 from engine control206 is provided to a pulsing valve 185, a pulsing valve 187 and/or apulsing valve 189, respectively, of each fuel nozzle 68. It will benoted that pulsing valves 185, 187 and 189 are located within a pulsingvalve housing 191 (see FIG. 4). Among other various readings, signalsand measurements received by engine control 206, a signal 216 is alsoprovided thereto by at least one pressure sensor 218 located adjacent toouter liner 76 of combustor 26 (see FIG. 2). Pressure sensor 218 sensesa frequency and an amplitude for the pressure within combustion chamber62 and imparts this information to engine control 206 via signal 216.Pressure sensor 218 is capable of withstanding high temperaturesexperienced in combustion chamber 62. Accordingly, an exemplary pressuresensor is a diaphragm type of transducer, where the displacement of thediaphragm is proportional to the dynamic component of the input pressuresignal. While only one pressure sensor 218 is depicted in FIGS. 2 and 5,it is preferred that a plurality of pressures sensors 218 be equallyspaced circumferentially around outer liner 76 in order to detectdynamic pressure of combustion chamber 62 in a more localized region.Accordingly, only those mixer assemblies located adjacent a region ofcombustion chamber 62 experiencing dynamic instability are modulated

More specifically, FIG. 6 depicts a schematic diagram indicating theflow of fuel from fuel pump 202 to circuits 180, 182 and 183. It will beseen that fuel pump 202, which includes both a booster pump 220 and amain pump 222, receives fuel from an inlet 224. Fuel pump 202 sends fuelthrough a line 226 to a metering valve 228, where the pressure iscontrolled. In order to maintain a desired pressure for the fuelentering main pump 222, a bypass circuit 230 is in flow communicationwith line 226. Bypass circuit 230 includes a bypass line 232 with abypass valve 234 therein for controlling flow back to main pump 222 viaa bypass input line 236. It will also be noted that fuel nozzle control204 taps into line 226 upstream of metering valve 228 via line 238 sothat it receives a high pressure source to modulate.

Upon exiting metering valve 228, line 240 splits first into a fuelsupply line 242 that provides fuel to a fuel supply manifold 244, whichin turn supplies fuel to valve housing 174 of each fuel nozzle 68. Aline 246 also in flow communication with line 240 is connected to fuelnozzle control 202, which enables it to determine a differentialpressure control of the pressure control nozzle (DPCPFN) and a torquemotor current of the pressure control nozzle (TMCPFN). From thisinformation, a fuel signal circuit 248 from fuel nozzle control 202controls the activation of staging valves 184, 186 and 188. Morespecifically, fuel signal circuit 248 includes signal 210, alsounderstood herein to be a pressure control pressure off the fuel nozzle(PCPFN), to a fuel signal manifold 250, whereupon fuel signal manifold250 then provides a signal 252 to each valve housing 174. It will beappreciated that staging valves 184, 186 and 188 will generally beactivated according to signal 252 so that the desired amount of fuelprovided via fuel supply manifold 244 is passed to the respectivecircuit of pilot mixer 102 (i.e., first and second circuit 180 and 182)and main mixer 104 (i.e., third circuit 183).

A signal fuel return line 254 extends from each valve housing 174 so asto be in flow communication with fuel pump inlet 224. A sink line 256from fuel nozzle control also connects to signal fuel return line 254.

It will be further seen in FIG. 6 that fuel nozzle control 202 receivessignal 208 from engine control 206. Under certain specified conditions,signal 208 instructs fuel nozzle control 202 to alter the distributionof fuel to circuits 180, 182 and 183 by activating staging valves 184,186 and 188 in a different manner. This occurs when the amplitude of adynamic pressure instability is detected in combustion chamber 62 abovea predetermined level by one or more pressure sensors 218. While thispredetermined pressure amplitude level may vary or be conditioned uponother engine factors, it generally will be set at a level whereintegrity of the combustor hardware is maintained (e.g., approximately0.5 psi peak to peak).

Besides altering the fuel split between circuits 180, 182, and 183,engine control 206 may respond to such pressure instability by causingfuel to be pulsed through one or more of pulsing valves 185, 187 and/or189. Pulsing of fuel through secondary fuel injection ports 134 of pilotmixer 102 in at least one mixing assembly 100 located near theoccurrence of the dynamic pressure instability, via pulsing valve 187,is typically preferred. It has been found that pulsing the fuel with anamplitude and frequency opposite that of the pressure dynamic reducesthe pressure instability in that location of combustion chamber 62.Alternatively, pulsing of the fuel may be done at an amplitude andfrequency which is a sub-harmonic of the dynamic pressure on thecombustion chamber. Pulsing fuel in this way would be at a lowerbandwidth, which would reduce the stress on pulsing valve 187 andincrease the life thereof. By utilizing a closed loop system ofdetecting pressure instabilities and then offsetting them through thepulsing of fuel in this way, the problem is attacked continuously untilthe dynamic pressure instability is below the predetermined level.Although fuel could alternatively be pulsed through primary fuelinjector 110 of pilot mixer 102 via and/or fuel injection ports 128 ofmain mixer 104 to offset dynamic pressure instabilities in combustionchamber 62, such as the case when pilot mixer 102 does not includesecondary fuel injection ports 134, it will be appreciated that pulsingfuel flow to secondary injection ports 134 has a minimal effect on thefuel/air mixture within mixing assembly 100.

It has also been found that a frequency signal from pressure sensors 218within a specified range is indicative of an incipient lean blowoutcondition for combustor 26. This signal range is approximately 40 Hertzto approximately 50 Hertz and is able to predict the oncoming conditionas opposed to merely detecting it. Accordingly, an override signal 214is preferably provided by engine control 206 to valve housing 174 sothat additional fuel can be supplied to mixer assembly 100. Preferably,override signal 214 involves the activation of valve 186, wherebyadditional fuel is injected into pilot mixer 102 by means of secondaryfuel injection ports 134. The fuel split between pilot mixer 102 andmain mixer 104 may also be altered by increasing the amount of fuelprovided to primary fuel injector 110 (e.g., when pilot mixer 102 doesnot include secondary fuel injection ports 134).

Thus, it will be appreciated that modifying the fuel split between pilotmixer 102 and main mixer 104, and even between primary fuel injector 110and secondary fuel injection ports 134, effectively counters the dynamicpressure instabilities in combustion chamber 62 and an incipient leanblow out condition for combustor 26. Likewise, pulsing fuel in primaryfuel injector 110, secondary fuel injection ports 134, and/or fuelinjection ports 128 is effective for the same purposes.

It will also be understood that control system 200 is also effective forcontrolling the pressure dynamics in combustor 26 when actions thereinare initiated intentionally. For example, it may be desirable in certaininstances (e.g., to improve the mixing of fuel and air during fuel richconditions) to pulse the fuel provided to mixer 100. Such pulsing offuel in and of itself may create pressure dynamics which need to bemaintained within acceptable limits. Detection and control of suchpressure dynamics by means of pressure sensors 218 and engine control206 may cause the pulsing of fuel to be modified accordingly.

In conjunction with the physical embodiments of mixer assembly 100 andfuel flow control system 200, it will be understood from the flowdiagram in FIG. 7 that a method of actively controlling fuel flow tomixer assembly 100 is also presented. More specifically, such methodincludes the following steps: sensing dynamic pressure (frequency andamplitude) in combustion chamber 62 of combustor 26 via pressure sensors218 (box 260); providing signal 216 containing frequency and amplitudeinformation of such pressure to engine control 206 (box 262); and,determining whether the frequency component of signal 216 is within aspecified range indicative of incipient lean blow out (comparator box264). If the frequency component of the pressure signal 216 is withinsuch specified frequency range, then engine control 206 provides signal214 to valve housing 174 to override the current status of stagingvalves 184, 186 and 188 to inject additional fuel into pilot mixer 102(box 266). Afterward, the dynamic pressure in combustion chamber 62continues to be sensed as represented by a feedback loop 267 to box 260.

Should the frequency component of signal 216 not be within the specifiedfrequency range, then the next step in the process is determiningwhether an amplitude component of signal 216 is greater than thepredetermined level indicative of a dynamic instability (comparator box268). If this is found to be so, then engine control 206 provides signal212 to activate pulsing valve 187 (and/or pulsing valves 185 and 189)and thereby modulate pilot flow at a frequency and amplitude whichabates the dynamic instability (box 270). Thereafter, the dynamicpressure in combustion chamber 62 continues to be sensed as representedby a feedback loop 272. Should the amplitude component of signal 216 beless than the predetermined level, the system likewise returns tosensing the dynamic pressure in combustion chamber 62 as shown byfeedback 274 connecting to feedback loop 272.

FIG. 8 further illustrates a staging diagram for mixer assembly 100,whereby the relative amount of fuel provided to pilot mixer 102 and mainmixer 104 is provided for various points in the cycle of engine 10(i.e., to obtain certain temperature ranges for combustor 26). Becausepilot mixer 102 includes both primary fuel injector 110 and secondaryfuel injection ports 134, it has been found that engine 10 is able tooperate at an extended temperature range when only providing fuelthereto. This also enables fuel nozzle control 204 to eliminate aseparate fueling mode (i.e., 60% pilot mixer/40% main mixer) which hasbeen utilized previously. As seen in a bar 275 in FIG. 8, the firstfueling mode involves 100% of the fuel being provided to pilot mixer 102to obtain a combustor temperature range of approximately 200° F. toabout 800° F. Bar 275 further depicts that a first cross-hatched portion276 thereof is attributed to fuel being provided only to primary fuelinjector 110 (i.e., to obtain a combustor temperature range ofapproximately 200° F. to approximately 500° F.) and a secondcross-hatched portion 278 represents fuel being provided to both primaryfuel injector 110 and secondary fuel injection ports 134 (to obtain acombustor temperature range of approximately 500° F. to approximately800° F.). This first stage is considered to be the range of normaloperation for optimum performance of combustor 26 when pilot mixer 102only is fueled. Thus, this first fueling mode is typically used foridle, taxi and approach portions of engine operation.

It has been found that a fuel pump limit 281 for the first fueling modeis reached at approximately 800° F. Accordingly, a second fueling modeinvolving some distribution of fuel between pilot mixer 102 and mainmixer 104 is required. As indicated by bar 280, the preferred fuelingmode for achieving combustor temperatures at approximately 800° F. isfor about 20% of the fuel to be provided to pilot mixer 102 and about80% of the fuel to be provided to main mixer 104. Utilization of thisfueling mode prior to this temperature point (as represented by blankportion 282 of bar 280) is possible without adverse outcome, but notconsidered to provide optimum performance of combustor 26. It will alsobe seen that a lean blow out limit 283 for this fueling mode is at acombustor temperature of approximately 525° F. The second fueling modeis used during a combustor temperature range of approximately 800° F. toapproximately 950° F., which is depicted by cross-hatched portion 284 ofbar 280. This second fueling mode is then utilized during the climb andcruise portions of engine operation.

It is then seen from bar 286 that a third fueling mode is preferred whenthe temperature of the combustor inlet air reaches approximately 950° F.It is preferred that the third fueling mode preferably includeapproximately 8% of the fuel being provided to pilot mixer 102 andapproximately 92% of the fuel being provided to main mixer 104. Thisthird temperature stage is represented by cross-hatched portion 288 ofbar 286 and involves a combustor temperature range of approximately 950°F. to approximately 1100° F. Utilization of this third fueling modeprior to this temperature point is possible without adverse outcome (seeblank portion 290 of bar 286), but not considered to provide optimumperformance of combustor 26. It will be noted, however, that a lean blowout limit 292 does exist at approximately 700° F. It will also be seenthat the second fueling mode (i.e., 20% pilot mixer/80% main mixer)could be utilized during this combustor temperature range (approximately950° F. to approximately 1100° F.) without adverse outcome (see blankportion 291 of bar 280), but it has not been found to provide optimumperformance of combustor 26. Implementation of the third fueling mode istypically done when the greatest thrust is required from engine 10, suchas during the take-off portion of operation. It will then be seen that afuel pump limit 294 is reached for the third fueling mode (i.e., 8%pilot mixer/92% main mixer) at approximately 1100° F.

Although particular embodiments of the present invention have beenillustrated and described, it will be apparent to those skilled in theart that various changes and modifications can be made without departingfrom the spirit of the present invention. For example, it will beunderstood that the method and apparatus of the present invention may beutilized with mixers having different configurations. While the mixershown herein has a pilot mixer with both a primary fuel injector andsecondary fuel injection ports, it may also be one where only theprimary fuel injector is provided. Accordingly, it is intended toencompass within the appended claims all such changes and modificationthat fall within the scope of the present invention.

1-35. (canceled)
 36. A method of actively controlling fuel flow from afuel pump to a mixer assembly of a gas turbine engine combustor, saidmixer assembly including a pilot mixer and a main mixer, wherein saidpilot mixer further includes an annular pilot housing having a hollowinterior and a primary fuel injector mounted in said pilot housing andadapted for dispensing droplets of fuel to said hollow interior of saidpilot housing, said method comprising the following steps: (a)continuously sensing dynamic pressure in a combustion chamber of saidcombustor; (b) determining whether an amplitude of the sensed dynamicpressure in said combustion chamber is greater than a predeterminedamount; and, (c) signaling a fuel nozzle to provide fuel in a specifiedmanner to said pilot mixer when said pressure amplitude is greater thansaid predetermined amount.
 37. The method of claim 36, furthercomprising the step of providing increased fuel flow to said primaryfuel injector of said pilot mixer.
 38. The method of claim 36, furthercomprising the step of pulsing fuel through said primary fuel injectorof said pilot mixer.
 39. The method of claim 36, said pilot mixerfurther comprising a plurality of secondary fuel injection ports forintroducing fuel into said hollow interior of said pilot housing,further comprising the step of providing increased fuel flow to saidsecondary fuel injection ports of said pilot mixer.
 40. The method ofclaim 36, said pilot mixer further comprising a plurality of secondaryfuel injection ports for introducing fuel into said hollow interior ofsaid pilot housing, further comprising the step of pulsing fuel throughsaid secondary fuel injection ports of said pilot mixer.
 41. The methodof claim 36, wherein said predetermined amount is a pressure amplitudeof at least approximately 0.5 psi peak to peak.
 42. The method of claim36, further comprising the step of continuously providing fuel to saidprimary fuel injector.
 43. The method of claim 36, further comprisingthe step of modifying a rate of fuel provided to said primary fuelinjector.
 44. A method of actively controlling fuel flow from a fuelpump to a mixer assembly of a gas turbine engine combustor, said mixerassembly including a pilot mixer and a main mixer, wherein said pilotmixer further includes an annular pilot housing having a hollow interiorand a primary fuel injector mounted in said pilot housing and adaptedfor dispensing droplets of fuel to said hollow interior of said pilothousing, said method comprising the following steps: (a) continuouslysensing dynamic pressure in a combustion chamber of said combustor; (b)determining whether a frequency of the sensed dynamic pressure in saidcombustion chamber is within a predetermined range; and, (c) signaling afuel nozzle to provide fuel in a specified manner to said pilot mixerwhen said pressure frequency is within said predetermined range.
 45. Themethod of claim 44, further comprising the step of providing increasedfuel flow to said primary fuel injector.
 46. The method of claim 44,further comprising the step of pulsing fuel through said primary fuelinjector of said pilot mixer.
 47. The method of claim 44, said pilotmixer further comprising a plurality of secondary fuel injection portsfor introducing fuel into said hollow interior of said pilot housing,further comprising the step of providing increased fuel flow to saidsecondary fuel injection ports of said pilot mixer.
 48. The method ofclaim 44, said pilot mixer further comprising a plurality of secondaryfuel injection ports for introducing fuel into said hollow interior ofsaid pilot housing, further comprising the step of pulsing fuel throughsaid secondary fuel injection ports of said pilot mixer.
 49. The methodof claim 44, wherein said predetermined frequency range is approximately40 Hertz to approximately 50 Hertz.
 50. The method of claim 44, whereinpressure frequency within said predetermined range is indicative ofincipient lean blow out in said combustion chamber.
 51. The method ofclaim 44, further comprising the step of continuously providing fuel tosaid primary fuel injector.
 52. The method of claim 44, furthercomprising the step of modifying a rate of fuel provided to said primaryfuel injector.
 53. A method of actively controlling fuel flow from afuel pump to a mixer assembly of a gas turbine engine combustor during aplurality of operational stages, said mixer assembly including a pilotmixer and a main mixer, wherein said pilot mixer further includes anannular pilot housing having a hollow interior, a primary fuel injectormounted in said pilot housing and adapted for dispensing droplets offuel to said hollow interior of said pilot housing, and a plurality ofsecondary fuel injection ports for introducing fuel into said hollowinterior of said pilot housing, said method comprising the followingsteps: (a) supplying fuel only to said primary fuel injector and saidsecondary fuel injection ports of said pilot mixer during a firstfueling mode; (b) supplying fuel to said pilot mixer and said main mixerin a first specified amount during a second fueling mode; and, (c)supplying fuel to said pilot mixer and said main mixer in a secondspecified amount during a third fueling mode.
 54. The method of claim53, wherein said first fueling mode is used to obtain a temperaturerange of approximately 200° F. to approximately 825° F. in saidcombustor.
 55. The method of claim 53, wherein duration of said firstfueling mode is determined by a limit of a pump supplying fuel to saidmixer assembly.
 56. The method of claim 53, wherein said second fuelingmode is used to obtain a temperature range of approximately 825° F. toapproximately 950° F. in said combustor.
 57. The method of claim 53,wherein approximately 80% of said fuel is supplied to said main mixerand approximately 20% of said fuel is supplied to said pilot mixerduring said second fueling mode.
 58. The method of claim 53, whereinsaid second fueling mode is limited by a lean blow out limit at a firstend and a limit of a pump supplying fuel to said mixer assembly at asecond end.
 59. The method of claim 53, wherein said third fueling modeis used to obtain a temperature range of approximately 950° F. toapproximately 1125° F. in said combustor.
 60. The method of claim 53,wherein approximately 92% of said fuel is provided to said main mixerand approximately 8% of said fuel is provided to said pilot mixer duringsaid third fueling mode.
 61. The method of claim 53, wherein said thirdfueling mode is limited by a lean blow out limit at a first end and alimit of a pump supplying fuel to said mixer assembly at a second end.62. The method of claim 53, wherein said first fueling mode is usedduring idle, taxi and approach portions of an operating envelope forsaid gas turbine engine.
 63. The method of claim 53, wherein said secondfueling mode is used during climb and cruise portions of an operatingenvelope for said gas turbine engine.
 64. The method of claim 53,wherein said third fueling mode is used during a take-off portion of anoperating envelope for said gas turbine engine.